text:
Why do jet aircraft need chocks when they have parking brakes?
I work on the ramp servicing small turboprop aircraft (ATR 72 and Q300s) and recently I’ve noticed that they don’t really require us to indicate when we’ve chocked the aircraft, however the larger jets do, as there is something they have to do involving the parking brake. So my question is what is so different about larger aircraft that they need to rely on chocks being in place? Aren’t all hydraulic brakes the same?

I assume it has more to do with turnaround timing then type.

Not sure how the parking brakes on a airplane works but I believe they can not be engaged immediately after landing, when still hot. And extra safety doesn't hurt. Cars are also choked before a mechanic would work on a car, even if the both the P-gear and parking brake are engaged.

@user3528438 While the parking brake shouldn't be engaged when the brakes are extremely hot, after a normal landing the parking brake is used when arriving at the gate. It's kept on at least until the engines have shut down and the ground crew has chocked the aircraft.

Reading the answers, the right questions seems to be: why to small turboprop aircraft not require the chocks?

@jcaron They are also chocked when parked longer, but with short turnaround times hydraulic pressure loss is negligible. The brakes are also rarely as hot as on jet aircraft because propellers in beta or reverse do most of the stopping.

...in aviation everything is about "if that fails then this" is the backup...

The reason you don't need to indicate this for smaller aircraft is that the pilot can see them before getting in.

First, chocks are an added level of safety. At many airports it is a local restriction that the passenger bridge cannot be connected before the chocks are in place.
During landing and long in-taxis the brakes sometimes get quite hot. Having chocks in place lets pilots to release parking brakes while on stands. This helps brakes to cool down as air is able to flow between hot brake discs. The ability of brakes to absorb energy is dependant of their initial temperature so it is often required to cool the brakes before the next departure in anticipation for a potential rejected takeoff at high speed.
Lastly, all(?) transport category aircraft have hydraulic brakes. During longer turn-arounds like overnight stops brakes tend to slowly lose their pressure, leaving the chocks the sole medium keeping the aircraft in place.

787 and A220 have electromechanical brakes, which do not have the slowly-lose-pressure failure mode of hydraulic parking brakes.

Another advantage for chocks is that an observer outside the aircraft can visually confirm that the aircraft is chocked.

I was a crew chief on C5 and C141 aircraft in Altus, Oklahoma. Everything that was on the flightline had chocks set. In Oklahoma sometimes the flightline would be solid ice with the wind blowing at a million miles an hour. The chocks would help prevent the plane (truck, generator, whatever) from sliding around on the ice. You wouldn't think something like a C141 could be pushed around by wind but they sure were.
Chocks are also like belts and suspenders - where if somehow the parking brake is released, you still have the chocks preventing your plane full of fuel, liquid oxygen, and atomic weapons from rolling into another plane full of fuel, liquid oxygen, and soldiers.
The air force tended to do everything the exact same way every time no matter what the circumstances - that way you get in the habit of "the one right way", and no matter where you happen to be in the world, or how stressed you are, you know the way you've been trained will work.

Were tiedowns available for smaller equipment?

i can't remember but i'm going to say no. at least not for the planes.  we had metal rings in divots in the flightline that were used for grounding and securing things like jacks and generators.

Hydraulic bleed-down can be a problem with keeping a parking brake set.
Also, if parking brakes are set, a FBO can’t reposition a plane without a pilot disengaging the brakes. That’s why jets will have a card in the windshield that says “BRAKES OFF”.

Because of the hydraulic pressure drops as the parking brake applied due to the normal internal leak of brake system so in some aircraft like (737 ) the hydraulic pressure can be stand only for 8 hrs if parking applied so we need chocks to keep aircraft in safe parking position for long times .

My tangentially related training in aviation maintenance leads me to believe smaller aircraft probably have a separate system involved in setting a parking brake.  Something like a cable that has very little variance over time and so could hold the aircraft still for very long periods of time as set.
I say this because most aircraft operate on a, 'primary', 'secondary', 'tertiary' system mentality.  Which is to say, there's always a primary system for any core functionality, and then a secondary system you can fall back on, and then a tertiary system to use if both primary and secondary have failed.  Tertiary isn't always available, depending on implementation, but there will always be primary and secondary.
For larger aircraft, there is no cable for brakes.  It's hydraulics only, as a cable system simply wouldn't be enough to stop the aircraft, making such a cable pointless extra weight.  What you'd have instead is a hydraulic primary system driven by the jet engine hydraulic pumps, a secondary electric hydraulic pump provided with power via the APU, and then chocks as the tertiary(obviously only when stationary on the ground, but also the only long term solution for parking).
Smaller aircraft probably use a hydraulic primary, with a cable secondary(which can be used as a reliable long term parking brake on the ground), and chocks become the largely unnecessary tertiary option.
So it becomes important to note whenever chocks are installed on a large aircraft, but not really consequential for small aircraft.

I'm curious if there are any common GA planes that have a cable brake backup to a hydraulic system. I can't think of any.


text:
A question regarding airfoil parameters
If a hypothetical aircraft, which is a tailless straight flying-wing with no fuselage, has the following characteristics:

moderate forward sweep
an aspect ratio = 9
a dihedral angle of 3 degrees
a slight tapering of 5 degrees on each edge
wing twist (10 degrees of total incidence difference between centreline and wing tip)
positively cambered airfoil at centreline, slightly less at the root, minimally positive camber at the inboard section, symmetrical outboard section, negative camber from there towards the wing tips
thickness 10% at the tips, 15% at the root and 20% at the centerline
a low wing loading, i.e. 50 kg/m2
slightly forward CG

And if the required stability in yaw (proverse yaw) is considered to be ensured through sufficient wing twist and symmetrical airfoil/negative camber at the tips,
what kind of airfoil characteristics - or even better - what specific airfoil examples (or at least, a specific airfoil category) would provide both low induced drag and ample lift ability at low speeds of 50KT, but still have low resistance at higher speeds i.e. at 200 KT?
Specifically, how much camber and reflex camber change (in ratio to symmetrical) should be sought for each of the different segments of the wing centreline, root, outboard section and wingtip?
Or, would it be sufficient if the whole planform has an identitical airfoil shape (scaled for chord length due to taper)?
What would be the Centre of Pressure positions (in chord percentage) of such a wing at slow flight or in high-speed cruise?
——
I know it may be a fairly wide question with a myriad of possible answers but I would be grateful to see some examples, so I can have some notion on what the camber/reflex camber for these wing characteristics should be, and to better understand via examples.
Thanks

You should get this book: https://marskeaircraft.com/new-book-the-wing-%26-i  Marske uses reflexed airfoils with modest FORWARD sweep and a large center rudder and is probably the most successful designer of flying wing gliders,

Thanks for the comment - I was considering forward sweep as I saw some examples similar to what you posted - regarding the rudder, it seems like a solid way to ensure the handling is sufficient yet the numerous examples of proverse yaw by wing washout / tip negative camber etc. are intriguing - I would very much like to know if the reduction in drag by eliminating the rudder would be perhaps not such a big benefit because of additional drag due to the negative (or less lift) load at the tips by using a high wing twist difference

The main benefit of forward sweep is elimination of spanwise flow problems endemic to aft swept wings.  The Marske FWs have impeccable stall behaviour bc the inboard elevator is incapable of achieving stall AOA.Also check out Rol Klingberg's youtube channel.  Here is one vid where he talks about spin issues with swept flying wings due to tip stall https://www.youtube.com/watch?v=oVaQMRXh2SQ&list=PLNp_btR7M1GUUJF0TbaTVASomkf3mXDRt&index=2.  He is designing a swept flying wing ultralight glider and has a unique one-way tip rudder system (he calls "auto yaw")  that greatly enhances yaw stability.

I saw the klingberg wing on youtube a couple of weeks ago, and the video of his glider in the 80s, i will this video now. The forward sweep is very interesting but i need to understand better the negative aspects from it, other than bending moments; as i read, it seems that fwd swept wings need upward wing twist at the tips

I don't think the Marske wings use wash-in which would tend to promote tip stall which promotes spins.  Marske exploits a phenomenon where the initial flow separation at the TE renders the wing root elevator ineffective before the stall deepens, which prevents any part of the wing from being brought fully to stall AOA.  The story of the Genesis is quite interesting because Marske's partners more or less forced him to add a small horizontal surface at the top of the fin.  He insisted it was totally unnecessary but added it when the partners threatened to kibosh the program if he didn't give in.

For a wing to be stable in its own right, its aerodynamic centre must remain static with changes in angle of attack. The usual approach is to sweep the wing and wash-out the outer section. Because your wing is unswept, you have only two choices in creating a stable aerofoil:

A symmetrical aerofoil, i.e. with zero camber.
A cambered aerofoil with reflex camber towards the trailing edge.

The first of these has poor low-speed characteristics and has really only seen use on aerobatic and early supersonic types.
The reflex type offers better low-speed characteristics, thanks to its positively cambered leading edge. RAF 34 provides a good example; it was developed in the 1920s when RAF 33, with even greater reflex, was found to have high drag. RAF 34 had much better performance above 200 kt and was used on several tailless types (usually swept and washed-out as well) such as the Westland Pterodactyls. It would be a good starting-point for your search.
But be aware that your tapered "plank" wing will have abysmal handling characteristics and a vicious, unrecoverable stall. Washout can help a little here, and the CG position will depend on this, as well of course as on your chosen aerofoil. Maybe using it as a hang-glider might be tenable, as the pendulum stability will mask the worst handling issues, but I would not want to take it up in case it stalled!
Addendum
Some remarks have touched on the NASA Prandtl-D experiments led by Albion Bowers, which owe much to Prandtl's "bell-shaped" lift distribution and Reimar Horten's exploitation of it in his flying wings. Prandtl's analysis applied principally to the straight wing and was not concerned with stability or yaw. By contrast the Horten/NASA implementation is applied exclusively to tailless swept flying wings, where stability and proverse yaw in the turn are essential to docile handling. In this latter context, many subtle variations of aerofoil across the span have been tried over the years. JW Dunne anticipated it all, and more, intuitively back in the pioneer era - the Science Museum archive contains boxfuls of his research models - but he was no mathematician. The precise relationships between the various design features and resultant flying characteristics still remain somewhat clouded and contentious, and Dunne's work studiously ignored, so some aspects of the original question here remain unanswerable. But one thing everybody I have mentioned has agreed on; if you do not sweep your wing then you are making life much harder for yourself.

Thanks for the answer- can you comment on how would forward sweep with RAF34 affect the AC? What about plank wing as I mentioned, but with RAF34 and a sizeable central rudder (as commented above by John K on the Marske glider)?  can you please comment also on this airfoil: http://airfoiltools.com/airfoil/details?airfoil=prandtl-d-root-ns ; it is the Prandtl-wing by Al Bowers, with proverse yaw. However, that aircraft has prominent sweep; there is this example: https://youtu.be/UsopLXgzpeM ; a version of the Prandtl-wing, no sweep, fairly delta-shaped with leading edge taper.

Also, in an answer on a question regarding wing twist (https://aviation.stackexchange.com/a/24470/53529) user Peter Kampf stated that forward swept wings can do away with twist by airfoil camber / having negative camber at the tips (if I understood correctly). Please let me know thoughts on this (I will edit question accordingly)

First query: A central fin on an unswept wing is useless, while the control surface acts more like an aileron than a rudder. Bowers' aerofoil and its AoA both vary significantly across his wing, the root is just one extreme. Unlike RAF 34 it was designed solely as the root of a swept wing. If you are chasing proverse yaw, then wing sweep plus washout plus tip anhedral (as per the Dunne D.7 monoplane https://en.wikipedia.org/wiki/Dunne_D.7 ) and Frise ailerons are your friends. Delta is equivalent to tapered and swept, see the Lippisch Deltas https://en.wikipedia.org/wiki/RRG_Delta_I.

Second query: I did not see forward sweep mentioned by Peter Kampf, perhaps I missed it; anyway, the forward swept wing needs upwards twist at the tips not downwards. Also, there are two kinds of twist, aerodynamic and physical. He seems to be talking of the aerodynamic type, which is a lot more subtle.

Perhaps I misunderstood when I read it, he wrote: “a positively swept wing can go away without twist if the airfoil is changed along the span from positive camber at the root to negative camber at the tip”.

Thank you for your feedback on my first two comments I need to think about these considerations.

"positively swept" means swept rearwards in the usual way. However Dunne pointed out in 1913 that negative camber at the tip (the Zanonia or crow wing) is directionally unstable. He gave the tips increased positive camber on his tailless swept wing (the seagull wing) for the world's first certified stable aeroplane, as witnessed by Orville Wright himself in 1910. To make the negatively-cambered tip type directionally stable, Jose Weiss, Handley Page and Igo Etrich all had to add tails. The physics of flight has not changed since then, though a shameful amount has been forgotten.

Ok so in that case, forward (negative) swept wings need a positive washout at the tips and swept-back (positive sweep) wings need a negative washout or a negative camber (although which is unstable). However as I re-read this one thing crops up to me as unclear, as you say that the wing must be designed in a way to maintain the aerodynamic centre “static”, yet, as far as I have seen the AC is a geometrical point which is constant in chord location with any AoA change?

Thx for the addendum it pretty much sums up the whole frame of what I didn’t see clearly, the interplay of sweep, twist and camber, yaw stability etc. Also, regarding my previous comment, I guess you mean that for a wing to be stable the pitching moments around the aerodynamic centre (the neutral point) should, at all angles below the critical AoA, result in the aircraft returning to the desired flight path?

The aerodynamic centre is defined as the point around which changes in aerodynamic pitching moments resolve (with changing AoA). The neutral point merely marks the CG position for neutral stability. For steady flight the centre of pressure must coincide with the CG, and it is around the CG that the net pitching moment must return the plane to level flight, by moving the CP appropriately.

Yes but from what I gather, the AC is also deemed an airfoil’s “neutral point” as, the whole aircraft’s NP is actually a location at which, if the aircraft rotates around it, the pitching moments are resolved and it is neutral in stability. Since aircraft rotate around the CG, only when CG is at neutral point will this occur. In the airfoil, the same thing implies, yet the wing doesn’t fly without the rest of the aircraft so.. airfoil AC is used because the whole aircraft has a CG which is not the same as the wing’s CG. So, AC is used, but it is basically the airfoil’s own ‘neutral point’

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text:
static thrust vs. prop RPM data?
edit: question was too long, and is actually simple, so I removed unneccesary description
Background:
There are many questions here about:

propeller efficiency
changes to thrust when it is dynamic
calculating the prop thrust for a given power
etc.

(Equation binding prop thrust, velocity and power)
(Units for prop thrust equations)
(Calculating thrust and required prop size for given engine power)
(Software for prop thrust and required power)
(Prop thrust equation taking into account number of blades)
(Prop efficiency of modern light sport aircraft)
(Correct formula for prop efficiency)
(Engine power vs. prop thrust for a 3-blade 14" prop)
But I would like to find more examples of measured static thrust made by a propeller.
So, if possible, please post some figures or some link with any data on measured static thrust on an aircraft with a piston engine.
Thanks
(example in photo below)

Thrust per HP will be somewhat higher if using a constant speed prop, closer to 4+ lbs/hp static.

Excellent remark I will edit that in the question. thanks!

Thrust depends primarily on five factors: prop diameter, prop speed, prop pitch, airspeed and horsepower. Ignoring four to focus only on horsepower isn't going to work, which is why it is hard to find the simple chart you want. Other factors like solidity also start to matter depending on disc loading and altitude, which drives density of air. You are probably best off looking at multiple specific cases to see how thrust changes with variation of one or two factors, like your Cessna chart.

Dynamic thrust, ofcourse, is a different matter, and altitude/density/airspeed etc are important for those calculations. These can be derived once static thrust is known. However all this is irrelevant to this question, as here the aim is to find examples of all three parameters measured, not calculated. There is no ignoring any factors as there is no focus on HP. The focus is on looking at data where RPM + Thrust (static, ofcourse......... as only this can be measured unless the whole airplane is in the wind tunnel) + kW (or HP) which can be easily inferred from RPM because torque is known...

but thanks for the comment let me edit that into the question so it will be more clear what is sought

You didnt mention rule of thumb, or? No I do not expect to calculate dynamic thrust. This is not mentioned anywhere. And even less calculating it just by extrapolating from static thrust. Dynamic thrust requires , as you also mention, other parameters... this is just a question to see what (if any) examples can be found of real measurements of static thrust vs RPM with atleast the engine known - a graph already showing static thrust vs RPM vs engine

Fine, I was only saying that your rule of thumb may work for some certain instances but is too simple to expect a research paper to substantiate. I didn't suggest that the cases you examine should be calculated. Static thrust only fixes the airspeed input. There are still three other primary inputs that vary.

i don't understand what are you referring to... firstly it is not my own figure it's a rule of thumb by the general aviation community , as rough an estimate as it is. It approximates a simple ratio which is not easily found (as there are many engines and propellers etc). And the aim of the question is not finding research papers etc but rather simple measurements of thrust ( static ofcourse) but at different RPM settings, not only full throttle, etc..

i see what you mean now, from the other user's answer, in reality there isn't any point in using static thrust for comparison of engine power conversion to thrust, as it shows something very basic. And all that can be calculated from parameters of prop diameter, blade angle, prop efficiency, air parameters, prop L/D etc. So basically, better to pursue the calculations to view thrust at prop and power produced by engine for that thrust?

Let's first clear up some nomenclature about "power".
Power into the system, or throttle, or rate of fuel consumption is one meaning.  Power as Force × Velocity is another, also known as mV$^2$/s.
These can be linked as potential energy/second fuel burn = Thrust × Velocity = Drag × Velocity (at steady state or constant Velocity.
More simply put Thrust = Drag
So, you can work this problem theoretically through calculation of form and induced drag for a given airframe in a range of speeds.  But, it is still difficult to calculate oweing to effects of "prop blast", which can affect both lift and form drag.
There is a wealth of data out there for various airframe/engine/prop combinations, which through years of testing have shown what is optimal, but it is really like trying to hit a moving target to develop a "Rosetta Stone" formula because of drag, weight, and desired speed range at what altitude all factor in.
But static test data would not be where I would start, as the relationship of prop pitch, RPMs, and airspeed is crucial.  All your Cessna 150 data is showing is that the prop AoA on the bench seems to be more efficient as RPMs increase.
More meaningful data might be airspeed vs RPMs in flight.  Thrust could be estimated from drag calculations, but in reality how the plane performs at a given throttle setting is what counts.
An approximation of drag at various airspeeds could be achieved by plotting angle of descent in a power-off glide.

glide ratio × mass = approximate Thrust at that airspeed

Discounting effects of changes in the airstream from turning props,  this yields thrust requirement for level flight, which could then be compared to RPMs required for level flight at that airspeed.
POST SCRIPT
So, attaching two props with bicycle chains and sprockets to a Belphagor turbine might improve thrust efficiency.  I believe Kuznetsov knows this.

thanks for this answer. You hit the nail on the head with the glide ratio! It is the fundamental cause of why I posted this question. Knowing the D/L gives the thrust required for an airspeed, but the issue here is that it concerns a specified airframe.

So, in contrast to that, I am not looking to know what thrust is required, and then being able to calculate everything deduced from that (prop rpm, engine kW, etc). What I am very interested to see is, for example, what sort of thrust would be at 2000 RPM for an engine which provides 150 kgF of static thrust at 3000 RPM with an output of 60 kW...? So, regardless of airframe

Also, by "power" i mean the power output on the engine's main shaft, the torque it produces at a given RPM.

so,  by "engine power" I mean what it usually means (engine output at the prop shaft) and not the potential energy of the fuel etc

@ivanantuns well, you could work with the theoretical prop airfoil "lift" (thrust) at a given rpm.  Even if pitch was not variable, AoA for a given relative wind could yield a theoretical thrust, but, when in doubt, overdesign it and test it safely please.  Note that prop RPM is also a function of prop drag.  I would stay away from static data.

overdesigning and being conservative, absolutely good advice! Thanks

Engine torque = Prop Drag torque, so if you know the lift to drag ratio of the prop airfoil, viola!

So, would that be the parameter which basically "governs" the whole situation. Can it be said that the L/D ratio of the prop blade at a specified blade angle (AoA), at a given prop speed (RPM), would give us the prop drag torque/engine torque requirement (which in turn gives the engine power output) for that RPM and propeller L/D.... and then again by knowing the L/D of the prop blade we can calculate the lift done by the propeller? Thus getting both thrust and rpm / engine power?

Yes!  But I would compare these calculations to actual results. Big long slow wings are far more efficient than rapidly spinning prop blades for a given amount of drag.  Next might be to plug in some horsepower - Newton meters/ second torque values to see how much angular drag the prop is making, then go ahead and compare it to static thrust data.  I think the L/D ratios will be much less than a similar wing airfoil. Pursue it.

I see. So basically, both thrust and engine power required for that thrust at a given RPM can be calculated by taking all known parameters: prop diameter, L/D and AoA [relative windflow at speed], prop efficiency [I guess this prop efficiency factor is what you refer to as compared to more efficient big long slow wings?], RPM, air density/pressure/temperature, airflow... etc.?

Yes.  But blade interference is an issue with static testing, which is why "push" at a given airspeed is elusive.  L/D (thrust per horsepower in) for props would be a great field of study.  I know helicopters get more out of theirs for above discussed reasons.  Good luck with this!

in any case, the fundamental cause behind this question is just that - comparison of real-life examples... Interesting how people tend to answer a question which asks "what examples of measured static thrust are there?" with "it would be very complicated to use these examples for further calculation let me explain these complications and why you shouldn't do that"

@ivanantums anyone can tie an engine to a bench and measure static thrust.  It's just that the data has very little application to aircraft.  Maybe it will help for us to remember props are like spinning wings, and they pull with lift.  This is why AoA and relative wind are so important.

I see. Now that is very very clearly put! I realize now why I was confused with this... the torque required for the propeller drag combined with the RPM available from the motor will simply give out a power rating. But across an operative RPM range and across the operative airspeed range (relative wind) at various blade angles the thrust will vary - so I actually need to study propeller efficiency throughout those situations and compare with actual results!! Thank you for all the input and the perseverance to clear this up!!!!

p.s. , what does the "p.s." in your answer mean ? why do you link wright bros , belphagor and nk12 ? :)

@ivanantuns sorry I could not resist that one.  Really trying to show how props (especially counter rotating ones) can be optimized.  Turbine give plenty of power, but props give thrust, especially at lower speeds.  Higher speeds generally not as good, but that NK-12 is does well at 500+.

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text:
Why has flight UA106 such a strange predicted flight path?
I was looking at the predicted flight path for UA106 from IAD to MUC and noticed this odd pattern. Why would it do this? I understand that most flight paths are not straight and there are airways that most all flights follow, but this path makes two less than 90 degree turns. Wouldn't it be simpler to go straight to the Belgium point rather than go to Poland, then back to Belgium, then back again to Munich?

FlightAware is notorious for bad flight tracks.

@RonBeyer Good to know. What is a better site to compare against?

@DavidK You could try flightradar24.com

This is an erroneous flight plan transcription. The actual flight path is correct. Probably the longitude of the Shanwick OCA exit point.

@SMSvonderTann I would say the opposite. FR24 frequently 'interprets' missing parts of the ADS-B data and makes a bad track. Flightaware does the same, but at least tells you when it does it.

Thanks @mins. The actual path definitely makes more sense to me. I edited the question to make clear that this is a predicted flight path. Would it be better for me to delete this question, or leave it up and have someone answer that predicted paths (particularly from Flight Aware) are often incorrect?

To me it's an interesting question, that opens potentially to different topics: Flight plan (routes and fixes listed on the site), great circles, FIR/OCA, NATS, coordinates... There is actually a problem explained in the meantime by DeltaLima. BTW I see that, for some reason, FlightAware doesn't use all the elements of the flight plan to display the predicted route (as visible here -- also incomplete, due to an incomplete database).

I'm voting to close this question as off-topic because it's effectively asking about bugs in a 3rd party service.

@DavidK it isn't even a "predicted" flight path, it just an incorrect rendering of the filed flight plan.

As @mins suggested, it seems to be a wrong longitude of a waypoint in the Shannon Control area. I assume it is a hemisphere error (they used East instead of West).
If the longitude of the waypoint is mirrored with respect to the Greenwich meridian, the resulting path gives a more likely route to be flown:

Would that be considered turbulance, @mins ?

Adding this link to FlightPlanDatabase.com, with the flight plan got on the page referenced in the question. Obviously the problem was on FA side with OLGON. OLGON is actually within Shannon FIR, not Shanwick.

Oops. We forgot the "-" in the fmc and our flight to Boston ended up in Kyrgyzstan!

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text:
What is the most efficient design for an aircraft designed only to hover?
What could be the most efficient propulsion solution/design for an aircraft (weighting less than 25 kg) designed only to hover (at a height of 100 m) ? No horizontal flight intended other then recovering the hovering position in case of atmospheric disturbances. Is ducted fan an option?

I would say hanged on a rope

Mounted on a pedestal.

Who needs propulsion go for a Tethered balloon

Or untethered.

How do you avoid horizontal flight on windy days? Wind speeds at altitude can be quite high.

What aircraft? A little 500 gram quadrotor has very different needs than a 10,000kg UH-60.

A bit of more context (altitude, payload, endurance, ...) might help in giving you a better answer

Kites are really, really cheap and have infinite fuel efficiency.

The key here is disk loading ($T/A$).  For most efficient hover, you want as low a disk loading as possible.  Small thrust (but must equal weight) and a huge disk.
When throwing a gas to make thrust (propellers, jets, and rockets), thrust is
$T=\dot{m}\,\Delta V$
I.e. the mass flow times the amount you increase the velocity of the air.
$\Delta V = V_j-V_0$
The mass flow is:
$\dot{m}=\rho\,A\,V_j$
I.e. the product of the density ($\rho$), the area of the jet ($A$), and the velocity of the jet ($V_j$).
For an equal amount of thrust, you can either accelerate a lot of air (large $\dot{m}$) or you can accelerate the air a lot (large $\Delta V$).
As it turns out, it is always most efficient to accelerate a large amount of air as little as possible (large $\dot{m}$, small $\Delta V$).
In hover, $V_0$ is zero, so, to keep $\Delta V$ small, you don't want a huge $V_j$.  So to maximize $\dot{m}$, you must maximize $A$.
So, we see that disk loading $T/A$ is the important parameter.

This is the most efficient type of aircraft designed only to hover. Tethered, sealed helium balloon. Unlike a hot air balloon, it doesn't require any propulsion, and altitude control is possible by inflatable air bags or tether tension. Some helium still leaks through the seams and needs to be topped up twice a year.

Anything else is comically wasteful for the purpose, including a helicopter.

If this is for a station-holding drone, for minium power you probably would go with a coaxial helicopter with the longest rotor blades you can live with, to get the highest L/D from the rotor (imagine two 40:1 sailplanes flying in a circle, vs two stubby wing 8:1 airplanes flying in a circle, both with the same weight and wing area; the gliders will need way less power).
At some point really long rotor blades become a problem for other reasons, and you'll have to find a sweet spot that provides decent control with minimum power.


text:
Which is the smallest available widebody on the market?
I'm interested to find out which is the smallest (capacity wise) modern widebody aircraft in production/ operation.

Smallest in which measure(s)?

Because the number of rows of seats can vary within a type, I propose the 767.  Narrowest inside width (4.72 m), fewest seats per row in economy (2-3-2).
For instance, United's 767-300 "Version 3 (46/121)" with a small economy class section has 46+22+43+56 = only 167 seats.

This question is broad enough to try out the new table feature:

Smallest by
All time
In production

MTOW
Airbus A310-200 (144t)
Boeing 767-300F (186.9t)

Length
Airbus A310-200 (46.66m)
Boeing 767-300F (54.94m)

Outside Width
Boeing 767 (5.03m)
Boeing 767 (5.03m)

max. Payload
Airbus A310-200 (32.8t)
Boeing 787-8 (43.3t)

Exit Limit
Airbus A310-200 (275)
Boeing 787-8 (381)

Passengers*
Boeing 767-200 (214)
Boeing 787-8 (242)

Seats Across*
Boeing 767 (7)
Airbus A330neo (8)

* number of seats for typical 2-class layout, seats across for typical economy
(source: Wikipedia)
In summary, the Airbus A310 was the smallest wide-body ever produced by most measures. The smallest wide-body still in production today is the Boeing 767, but only the freighter variant is still produced. The smallest passenger wide-body in production is the Boeing 787 Dreamliner.

I upvoted the 'by which measure comment', you just answered for all measures. Nice! +1. And good use of the new table feature!

For number of passengers, Wikipedia's table is incomplete.  Seating varies within each type, and therein even within each airline.

@CamilleGoudeseune I know, that's why I listed exit limit first. It is a clearly defined number. For passengers I listed what Wikipedia listed as typical in a 2-class layout. Actual numbers can vary a lot of course.

Ah, that makes sense.  +1.


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Any ideas on what caused my engine failure?
On my second solo flight I came back to do a few landings, and after the first landing I applied full power for takeoff to which the engine responded by just shutting off on me. Not sure if my mixture was too lean, if I applied power too quickly (which I read can temporarily lean mixture too much to keep the engine running), or something else. My CFI wasn't sure what caused this either.
After an embarrassing talk with ATC I was able to get it started up again just fine, and the plane has ran well ever since, so I'm 99% sure I caused the incident. Just not sure exactly what.
I fly a 1976 Cessna 150M. Airport elevation is 5500 above sea level (hence why I leaned my mixture in the first place, as it's advised in the POH). Weather that day was around 0C but very arid.

Unless you were operating at a high altitude airport why was your mixture anything but full rich on approach/go-around?

@Dave Because that was what I was taught. I have never landed or seen my CFI land with mixture full rich, nor has he ever asked me to land this way.

Hmm... as Dave pointed out, mixture full rich should be a memory item...

Welcome to aviation.SE! It would probably help to add the aircraft type, engine, airport, weather conditions and any other details you think might be relevant. Even with that information it might not be possible to identify a reason definitively, of course.

Was your carb heat on? Aircraft carburetors seem to be able to make ice even at zero humidity.

@A.I.Breveleri Almost certainly, as that's what I always do before descending, but since it's been a week or so I can't directly remember turning it on. So in that respect I suppose it's possible that it may not have been.

Closely related:  https://aviation.stackexchange.com/questions/12891/why-must-the-throttle-be-moved-slowly/12911#12911

seems to me the only mistake you made was "I was able to get it started up again just fine, and the plane has ran well ever since" - I wouldn't touch an airplane that cut out on me for no reason.

@Dave wasn't Allihusk indeed operating at a high altitude airport? 5500 ft? seems to me it was.

@java-addict301that was a later update and was not included in the original question (hence my note)

@java-addict301 If the engine had backfired, started smoking, or otherwise behaved violently I would not have done that. In this instance it literally just smoothly shut off, as if I had cut off the mixture. With my (limited) knowledge on engines, I didn't see anything unsafe about simply starting up again.

The most likely cause of your engine cutout is what's known as a rich cut caused by coarse application of full throttle. Most carburetors have an accelerator pump which jets extra fuel when the throttle is pushed in quickly, this is to make sure the engine has enough fuel and prevent what's known as a lean cut. However, at high density altitude your mixture needs to be leaner than at sea level, so it can in some circumstances do the opposite and give your engine too much fuel to burn.
It's also possible your mixture was too lean, so advancing the throttle caused it to die on you. There's no way to know for certain.
In the future advance the throttle smoothly, it doesn't have to be super slow, just don't jam it in. 2-3 seconds from idle to full is about right (same goes the other way too). If it starts going gluk gluk gluk on you enrich the mixture a couple of turns and see if that resolves it. Don't forget on a touch and go you have the option to abort, don't be afraid to use it if there's issues. Remember that takeoffs are optional, landings are mandatory.
And remember that carb heat! Icing can happen in any conditions.

Yes, I had this happen (once - never say I don't learn from experience!).  Not on takeoff, but wanting to make a climb.  Luckly the prop was windmilling, so the engine caught after a couple of seconds.

That must have been an interesting few seconds @jamesqf!

I had the opposite happen back in the mid 70s flying a flying club Bowers Flybaby homebuilt that had a Cont C90 with no accelerator pump and no starter.  It just died when I added throttle a bit too abruptly on final and I rolled to a silent stop halfway down the runway. I arrived back at the ramp a little tired, to the surprised look of some of the locals, walking with the tail on my shoulder, pulling the thing along backwards like Jesus dragging the cross (a cross with handy rubber tires; the Flybaby was a small single seater and was light enough to do that).


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How often are turbine blades in jet engines replaced?
How often are the turbines in the hot section of the jet engines replaced due to deteriorated TBC? or just how often are they replaced? lets say for CFM56-7

Related: https://aviation.stackexchange.com/questions/71091/turbines-in-jet-engines

You can restrict your question to a specific model of engines as it may vary from engines to engines.

So this article, says that the average brand new CFM56-7 will go 18,000 cycles (i.e. flights) before its first shop visit.  You can't replace the turbine blade on-wing, only in the overhaul shop.  So we could say that the blades would last at least 18,000 flights, potentially longer.


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What do I look for when getting a PPL at a college or local school?
I am interested in getting a private pilots license and I know I can go to a near by airfield to get one. However I would like to know if anyone has received one through college and if it would be best just to wait until then? The cost at the airfield near me is around $12,000 but i have not found a cost for getting one in college.

You might want to mention in which region you live, or in which region(s) you would consider to acquire the license.

The Midwest more, specific Indiana. I know Purdue has programs like and Western Michigan does as well. Other than those two I've had a hard time finding colleges that provide private pilots licenses.

Find the lowest cost flight school you can.  Part 61 or part 141 is irrelevant.  You will pay considerably more for a certificate at a college but the quality of instruction could be lower.  College instructors are typically low time instructors building time for airlines.  A local school may have more experienced instructors who teach because they enjoy it.  Just a thought

I believe Vincennes also has a flight program, and I think Rose Hulman does as well.

@ThundaChicken76 "What is best" questions will always get closed. However, if you were to rephrase it "What should I consider when comparing a & b" you'll likely get the question reopened.

There are 2 major drawbacks to getting a PPL through a college program.
1) Admissions
In the comments you mentioned Purdue University. An excellent school, but they have very high academic standards (my son was interested in going there - didn't get accepted). There may be some variation for admission to the flight school program vs other degree tracks, but you'll still have to get admitted to the university of choice.
2) Cost
When getting your PPL in conjunction with a college degree, you're going to be taking a whole lot of general education courses that don't apply to flying. To go back to your Purdue example, they will not want you parading yourself as a proud Purdue alumn, yet not have the basic education levels that people would expect from a Purdue grad (IU v Purdue sparing aside). That means you're going to be taking writing and basic math and humanities classes, none of which have anything to do with getting your PPL. Each one of those credit hours will cost you. As will your flight hours. A school like Purdue is a prestigious, Big 10 school and tuition there ain't cheap. Purdue (again, mentioned in the comments and I have some personal experience) will run somewhere in the neighborhood of $30,000 per year for in-state tuition. I don't know the length of their flight school program, but you could do 2.5 trips through the flight school you mentioned for that kind of money.
If you go the straight up flight school route, you're going to be paying for flying lessons and flight time and that's it. Not having gone to flight school, I don't know what the admissions criteria are, but I'd venture it's not much more than having a medical certificate and the cash to start paying. If you run out of cash, you can put your training on hold for a while.
Of course, if you're going to school to get a degree and you want to make getting your PPL part of the process, not the sole reason for being there, that's a different story as far as classes go, but you're still paying a lot for the privilege of taking flight lessons through a university.

@ZeissIkon I was gettin' there! I wanted to have an answer in before the question got closed...

The usual advice is not to answer "bad" (likely to be closed) questions -- if you know it's going to be closed, rushing in an answer isn't desirable.  Ideally, if a question is closed as OT, opinion based, or unclear, it shouldn't have answers when it closes.

@ZeissIkon that's quite true, and I've said that myself to others... I think this could be quite salvageable if it were rephrased "what do I look for" as opposed to "which is best". I'll go make that suggestion.

You go to a college because it's part of a career track and as such will be filled with all sorts of courses that are professionally useful but not directly related to learning to fly airplanes.  And it won't be any cheaper.  If you just want to learn to fly for recreation, find a local school with the best instructors and equipment.
Probably the best overall and cheapest way to do a Private it is to find 2 partners and buy a trainer (2 partners gives you the optimal dilution of fixed ownership costs), and hire an instructor to train the 3 of you.
And if you really want to learn to fly, properly, start with glider training first.

Thanks for the response, and I am planing on going to college for aviation technology/aerospace engineering. I just was curious on which would be more beneficial college or local school for PPL.


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Placing bomb bays on top in order to avoid radar detection
Considering that opening the bomb bay on a stealth aircraft increases the chances of the plane being detected by radars, why no one is placing the bomb bays on top of the airplane? At least for jet fighters like F-22 or F-35.
Edit: Asking this from a theoretical point of view. I am not asking for any specifics. Just if done, would it bring any advantage? And if yes, what would be the most reasonable way to do it. If it is not possible, why not? It wouldn't be the craziest idea to implement or at least to try.

Because that would make it pretty difficult to drop the bombs?

When you roll-upside down to drop the bombs you lose all the advantage.  Actually its even worse than that because the airplane is designed to minimize the radar reflections when the aircraft is upright, not inverted.

Two relatively naive answers. An F-117A was shot down by Yugoslavia with old tech simply because the bomb bay was open long enough. The same would affect F-22 or F-35 fighters I imagine. There would be various ways of adapting a top bomb bay by creating some sort of catapult. And I didn't mean to imply that rolling the plane over would be part of the solution. But opening the bomb bay (or let's say missile bay for those incapable of abstract thinking) still is the most important factor exposing a fighter.

Considering the lack of detail in the question, the comments you got are quite reasonable. You should expand your question including all the details you though of, explain why you think it would be a good idea, and show that you have done a minimum of research.

Basically this is just discussing an idea. From a theoretical point of view. I am not asking for any specifics. Just if done, would it bring any advantage? And if yes, what would be the most reasonable way to do it.

Welcome to aviation.SE! We generally try to avoid questions that ask for opinions and discussions; the [tour] might be helpful if you're new here. You might get a better response if you can explain how the aircraft would release the bombs (eject them upwards? fly inverted to drop them?) and how that would be better than the existing technology. For example, if your idea is to fly inverted and then open the bay, how is that an improvement over just flying level and opening a bay?

And the munitions are supposed to depart from the aircraft how, exactly?
Currently, bombs are kicked off downward and sometimes a little bit sideways as well. There is a small shotgun size charge that pushes an ejector foot down, that propels the bomb away from the aircraft. This, of course, works in conjunction with gravity.
A bomb bay on top would have to have a LOT larger impetus to get the bomb to clear the aircraft.
Missiles are a different thing. They are self propelled, going forward. But still need some clearance from the aircraft body.
Fuselage mounted missiles such as an AIM-7 on the F-15 still need to be pushed away from the body before firing the missile motor.
There has been at least one aircraft with over-wing mounted missiles, the Jaguar GR3.
The weapons bay doors are only open for a short time.
Even as far back as the F-106, you could go from completely closed, open, launch all 4 missiles, close bay doors in about 4 seconds.
And from a personal experience aspect, loading bombs into a top opening bay would be a LOT harder.

Is it really that difficult to come up with a tech which would catapult the bombs or missiles? And I am talking mostly about jet fighters, not large bombers. Even 4 seconds is long enough to leave the plane exposed. Let's not forget the F-117A shot down by Yugoslavia with relatively old tech.

@TheFuriousSnail "Is it really that difficult to come up with a tech which would catapult the bombs or missiles?" looks like. Av.SE is not really the place to find new technologies, and if nobody has ever done it before, we'll argue that is probably not a good idea. If you have hard proof of the contrary, you're welcome to share. But if you're here just di discuss an idea, it might not be the right place.

The F-117 was shot down due to a dumbass at the stick. Flying the same route, at the same time of day, repeatedly. These things aren't invisible, just harder to detect. If you know when and where it is going to be, toss up a few missiles in the path. One might get lucky. And it did.

As for designing a mechanism to do this? I'll leave it to you to calculate the force needed to raise 500-1000kg 2 meters straight up, in 1 second. How much would all this extra catapult hardware weigh? This cuts into your range and/or payload.

@WPNSGuy I am aware stealth doesn't make you invisible. Claiming that the pilot was "a dumbass" is quite naive. His flight plan was most likely pre-approved by a bunch of people. It was a military or intelligence operation after all. According to Wikipedia the plane was detected exactly because the bomb bay was opened.

@TheFuriousSnail -  "However, that the NATO mission planners had complacently scheduled the stealth bombers on predictable, routine flight patterns. Worse, the Serbs had managed to break into NATO communications and could overhear conversations between U.S. fighters and the airborne radar planes directing them, allowing Dani to piece together a accurate picture of those routines."
https://nationalinterest.org/blog/reboot/thats-embarrassing-1999-yugoslavia-shot-down-f-117-nighthawk-166788

@TheFuriousSnail - The best aeronautical designers and engineers, from all over the planet. Chinese, Russian, French, British, German, American. Interestingly, no one has deployed a fighter aircraft with your design concept. I wonder why.


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Why are heavy flaps better than just a bigger wing?
Flaps increase lift during landing and T/O. But when retracted, they do nothing. The space needed to stow the common fowler flaps can't be used for anything else - fuel or structure. Extended flaps are thin, making them need more material to maintain stiffness. Moreover, they transfer their loads through the wing, instead of directly to the fuselage.
So, why do we use flaps instead of just scaling the clean wing up?

Is it because of concerns over cockpit visibility during landing? Can't we use video cameras do display the bottom during high AoA?
Is it because the extra lift - and drag - is unwanted? We can simply increase the cruising altitude without changing the cruise speed.
Is it because (for fowlers) the chord reduction really is that beneficial?  The larger wing volume for fuel leaves more space in the fuselage for everything else. And since the span does not change, induced drag does not increase.
Is it because slotted airfoil, as in many modern flaps, are not good for cruise? Surely we can add slots to the airfoil, and add mechanisms to seal them for cruise.

related: For a large commercial plane on landing, does the L/D ratio increase, decrease, or not change much?, this answer, and this answer (this last one also include speed polars for a specifi aircraft at different flaps settings).

Another closely related, if not a dupe: Why and when to use flaps?

"Close the slots for cruise" would take additional, heavy mechanisms that would need to be dragged along for the entire flight and take up space in the wing that couldn't be used for anything else. i.e. the same issues you complain about for flaps, therefore, even if lighter and requiring less space than flap-retract mechanisms, not that much of an improvement over the current situation of retractable flaps.

"even if lighter and requiring less space than flap-retract mechanisms" That is my point. Besides, we don't seem to be designing A/C's so much with simplicity as efficiency in mind.

This question completely overlooks the effect of flaps on the drag coefficient.

Not actually. You seem to be imagining a scenario where non-retractable flaps are used. Actually, what I'm saying is, the main foil obviously gives better lift/drag that main foil+deployed flaps. The idea is to scale this main foil up to the lift needed.

@ABJX -- see my answer

@ABJX, Why have a transmission in your car with multiple gears?  It adds weight and expense to the car.  Why not simply scale the gear ratio to provide a single setting?  Because it is possible to do this...

It takes a different design to fly faster than the speed of sound. Speed of sound decreases with altitude, up to about 90000 feet. This means that you will not be able to go higher and faster as the planes today already are close at the level they fly --higher up your imaginary plane would need to reduce speed.

why does this graph look like this then? https://www.researchgate.net/publication/274750587/figure/fig2/AS:615066709798913@1523654739040/Speed-of-sound-as-function-geometric-altitude.png

Your concerns about heavy flaps are well founded. The designers try to get away with as few high-lift devices as they can afford to. But not fewer!
If you observe the trend over the years, flaps became more complex with every new airliner generation, starting from simple split flaps in the 1930s to triple-slotted flaps on the Boeing 747 in the late Sixties. But then things reversed a bit. Now double-slotted flaps are standard and lighter versions of the same type (think A318 versus A321) get away with simpler flaps.
One reason is wing tank volume. In order to cross the Atlantic, the first generation of jets needed large wing tanks which were made possible by a high wing area. Simple, single-slotted flaps were sufficient for the desired landing speeds. With the much lower fuel consumption of high-bypass engines we now can afford smaller wings with less chord, but now the flaps have to make up for what is lost in area. High-lift devices are a major part of the aircraft development effort and a lot of work goes into reducing the complexity of flaps and slats. The variable-camber Krüger flaps of the 747 are great, but were never repeated on newer designs.
Even simpler wings would be possible if the cruising altitude were higher. But there is not much benefit from climbing above the tropopause (except for strategic bombers, but their development effectively ceased half a century ago), so that is where the installed thrust is optimized for. If you want to fly higher, you need larger and more expensive engines but gain little in cruise efficiency.
And to not retract the flaps is not an option. The larger area means that gusts can potentially put larger loads on the wing and the increased surface area would cause more friction drag. Reducing wing area saves fuel, even though the wing becomes heavier. On top, a heavily cambered wing would be completely unsuitable for transsonic flight.
Flaps have become much thinner in the last half century, and for good reasons. Yes, you need a complicated load path through flap rails and into the main wing, but that is where the stiffness is for carrying large loads. That should not be duplicated in order to keep structural mass low! The effort to reduce flap complexity has led to ever thinner flaps, and the development of transsonic airfoils with their high rear camber has allowed to put more camber on the flaps as well which improves their effectiveness. Note that the fairing of the flap tracks is used for area ruling and helps to limit the transsonic drag increase.

When flaps are retracted they do nothing, which is the whole point. The byproduct of lift is drag, a larger wing will create more lift, but more drag as well. More drag equals a slower cruising speed, or bigger engines to power past the drag along with higher fuel consumption. Flaps let airplanes cruise faster by getting out of the way.

Well, the flaps become dead weight. Also, as I said earlier, cruise alt could be raised

Great @ABJX, so you cruise slowly at a higher altitude, which it takes longer to climb to and descend from. As a regular passenger on commercial flights I don't call that a win. On a light aircraft a higher cruising altitude has no benefits, I can already cruise high enough to need an oxygen system.

You got me wrong there. When did I say it would have to be slower? If anything, you could go faster. Push the mach limit harder. Higher alt = less air = less air crashing uselessly into the fuselage. Anyway, back to the point. We scale the wing up for the landing condition. The increased volume allows for more of the plane's fuel to be stored in the wing, freeing up the fuselage. The increased lift is used to gain this higher altitude. The lack of dead weight flaps and the increased spanloading decreases weight.

Even if we don't speed up, we can reduce the fuselage fineness ratios to something more optimal for surface area. And the landing gear bays could move to the wings as well, freeing up even more space in the fuselage - or just allowing the belly to be mostly eliminated.

It's in the question.

You are assuming higher altitude is a good thing @ABJX. I spend my flying time in a light aircraft below 5000ft, I can't fly higher for the most part because of controlled airspace above. I wouldn't want to trade no flaps for a slower cruising speed.

Again. Higher alt is meant to allow the cruising speed to remain as fast as it is today. Hehe.

Uh, no @ABJX, it doesn't work that way. Higher drag at lower altitudes means higher drag at higher altitudes, you aren't going to achieve the same cruising speed just by going higher. Yes, air resistance is lower but you still have to move through it.

@ABJX Not to mention propulsion efficiency takes a hit at lower density.

@ABJX: "Higher alt = less air" = having to pressurize your airplane & turbocharge your engines, and spend more time climbing to cruise altitudes.  At which your jet engines may not even work...

Jetliners are pressurized to around 8,000-6,000 feet. They cruise at 30-40 thousand. The absolute value of the pressure difference from that to 60,000 feet doesn't look that big - that is, the additional cabin strengthening needed would not be that great. As for jet engines, that will solve itself, I assume.

@ABJX: I don't happen to own or fly a #%@! jetliner.  I'm not sure that my Cherokee would even climb with full flaps, or a wing equivalent to them.

Climb to cruise burns fuel.
Adding additional drag burns fuel.
Adding retractable mechanisms adds weight that burns fuel.
More drag, even at higher cruise altitudes, requires larger engines for the same cruise speed. Larger engines burn more fuel (despite increases in modern engine fuel efficiency).
Retracting high-lift, high drag devices reduces fuel burn (even though they add weight, thus drag which burns fuel).
Carrying the fuel necessary to carry that additional fuel burns fuel (several questions here about that, I invite you to look to see just how much it costs).
Increased fuel capacity reduces paying cargo (both boxed and self-loading) capacity.

Just like everything else in aircraft design, there's a trade off between the full-time lift/drag of a high-lift wing vs the weight/complexity of retractable high-lift devices on an otherwise low-lift wing.
Designers have decided that the reduction of fuel load in the wing and the additional weight and complexity of retractable flaps and slats to generate the lift necessary for safe and sane take off and landing speeds and runway length is a better bet than adding additional lift and drag, but additional fuel capacity, by designing a higher-lift wing.

"More drag, even at higher cruise altitudes" what?

@ABJX a fixed high-lift wing will generate more drag (even at a high altitude) than a flaps & slats wing. Guess that wasn't clear, edit coming.

@ABJX however, given any 2 aircraft at the same cruise altitude, the one with the higher Cd will require more power for the same speed, will it not? So, if you have a craft with a fixed, high-lift wing 5 miles ahead of, and at the same altitude as, a craft with a low-lift wing, but otherwise identical, the fixed high-lift winged craft will require more energy to maintain the cruise speed and stay ahead.

Aha, I get your problem. "a fixed high-lift wing" that is, "a wing with non-retractable slats and flaps" would certainly not work. But that is not what I was talking about. I was talking about eliminating the flaps and scaling the airfoil to compensate.

@ABJX Would "scaling the airfoil to compensate" not involve adding more lift which would be there, adding both lift and drag all the time, not be a "high life" wing, in comparison to a "standard" wing with retractable high-lift devices?

As so often, the best answer languishes at the bottom. +1.

High praise indeed, @PeterKämpf - Thank you!

@PeterKämpf Yeah, nice to see I am not alone.

The questioner seems to have noted that the basic wing with flaps retracted provides a high ratio of L/D (or Cl/Cd).  Where L denotes lift, Cl denotes lift coefficient, D denotes drag, etc.
We can certainly scale up the basic unflapped wing to provide as low a landing speed as we wish, although landing will be tricky due to the flat glide path.  Flaps help with landings by increasing the drag coefficient as well as the lift coefficient, making the glide path steeper.
The main problem with this approach is that for cruising flight, not only do we wish to achieve a high L/D ratio, we wish to achieve it at a high airspeed.  Lift is proportional to lift coefficient times airspeed squared, and in cruising flight, lift cannot be larger than weight.  If the wing is too large, it will be optimized to deliver its peak L/D ratio at a much lower airspeed than we wish to cruise at.  In such a case, at our intended cruise speed, if we increased the angle-of-attack to the max L/D angle-of-attack, the wing would be making way too much lift, and we'd pitch up into the start of a loop.  To keep the flight path level, we'd have to trim for an angle-of-attack far to the right side of the peak L/D ratio, as portrayed on the polar curve of L/D ratio versus airspeed.  In other words, we'd have to trim to an angle-of-attack much lower than the angle-of-attack that delivers the max L/D ratio.  We'd end up with more drag than we'd have if the wing were smaller.
The situation is not unlike that of a glider pilot wishing to achieve a flat glide at a high airspeed.  The glider gets a lower sink rate and a better glide ratio at that high airspeed when the wing loading is high than when the wing loading is low, because the wing may be operated at the angle-of-attack that yields the best L/D ratio, rather than at some much lower angle-of-attack.  So water ballast is carried.
In powered flight, the equations are a bit different, and simply adding weight to the aircraft never improves high-speed cruising performance.  But if the wing were designed to be large enough to give an acceptably low landing speed without using flaps, then scaling down that wing to a smaller size certainly would improve high-speed cruising performance.  And that's why it's worth carrying around the weight and internal volume of a complicated flap system-- because we can make the wing smaller, so that when the flaps are retracted, it is optimized to deliver its peak L/D ratio at a high airspeed.
The basic thrust of this answer remains the same regardless of whether we are trying to achieve a low landing speed by scaling up the wing in all dimensions, or by only increasing the chord.  In general, a high peak L/D ratio is associated with a high aspect ratio, and therefore a small wing chord.  However, if in cruising flight we know we must fly our scaled-up wing at some airspeed that is much higher than its maximum L/D airspeed, it's possible that we'll have a better L/D ratio at that high airspeed if we've scaled up the wing by expanding the chord alone than by expanding all dimensions.  Because the curve of L/D versus airspeed may be less "peaky" with the lower aspect ratio than with the higher aspect ratio.  But the better solution is to keep the high aspect ratio, and keep the wing small enough so that it can actually be flown at its maximum L/D ratio at the intended cruising speed.  Then we "scale up" the wing for landing by extending the flaps.
Of course, extending the flaps does much more than just "scale up" the wing.  At full extension, the designer's goal is to minimize the stall speed, so maximizing the lift coefficient is the priority.  He or she is free to choose a configuration that maximizes the lift coefficient, with no concern for minimizing the drag coefficient to maximize the L/D ratio.  As noted above, an increase in drag is actually helpful during final approach-- it is easier to guide the aircraft to the intended touchdown zone if the power-off glide path is not too flat.  (And no, despite the questioner's suggestion, we cannot eliminate the extra drag created by the extended flaps in cruising flight simply by "closing the slots".)  The flaps are designed purely to make the wing's lift coefficient as high as possible, while the airfoil of the clean wing is designed to optimize the ratio of L/D or Cl/Cd.  Therefore, to achieve the same low stall speed simply by scaling up the unflapped wing -- either chordwise, or in all dimensions -- would require a much greater increase in wing area than the area that is actually added by the deployed flaps.
Related --
(Q) Why would a glider have water ballast? If it is trying to stay aloft without an engine, wouldn't it be better to be as light as possible?
(A) For a large commercial plane on landing, does the L/D ratio increase, decrease, or not change much?
(A) For a large commercial plane on landing, does the L/D ratio increase, decrease, or not change much?
(A) Why and when to use flaps?

Ohhhhh sorry. I am learning not just aviation, but how to ask questions. " we cannot eliminate the extra drag created by the extended flaps in cruising flight simply by "closing the slots".)"  I'm gonna try to clarify more in the next edit. By the way, I see that even this answer ignored the option of higher cruising altitude.

@ABJX -- "By the way, I see that even this answer ignored the option of higher cruising altitude." -- yes, for once, I guess I figured that one of my answers had already gotten long enough.

@ABJX If you expect a higher cruising altitude to reduce drag then a higher cruising altitude will get a bigger drag reduction if you add flaps. Then you say "higher" to which the flapped wing can also go higher and have less drag. Continue until you reach service ceiling


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For jetliners, how is a swept wing justified over a thin wing?
For a given speed, a swept wing is a fatter alternative to a straight, thin wing. But why? Does not the sweep lengthen the wing and make it heavier? Is the swept wing chosen for the extra fuel tank volume?
Edit
Could you quantify the justification? Like, let's say we have this beam of depth x............

You add an edit to your question, I add an edit to my answer.

Yes, sweep makes a wing heavier and less efficient. But making the airfoil of a straight wing thinner runs into diminishing returns and also will drive up mass.
Since a thin wing needs to create the same lift as a thicker one, only the local speed increase from thickness can be tackled by making the airfoil thinner. The speed increase necessary for lift creation needs to be maintained, so the thickness contribution becomes a minor factor for the wings. However, a thinner wing requires a heavier spar to carry the same lift loads, so more lift needs to be created by the thinner wing to carry that extra weight. Now further thinning will drive up the speed increase from lift which will eat up the gains from thinning the wing.
If you look at parametric weight equations for airliner wings, the relative thickness at the wing root $\delta$ is part of them and generally occurs in a factor like
$$m_{Wing}\propto\frac{1}{\delta^{\,0.36}}$$
Now it is important to stress that this is valid for the range of thicknesses from which that relation had been derived, something between $\delta$ = 14% and 16%. If you drop below 14%, the exponent will most likely go up since the mass of the spar alone should be proportional to $\frac{1}{\delta}$ and becomes more dominant the thinner the wing becomes. But if root thickness is decreased from 15% to 14%, that means we stay within the limits of the statistical model, the wing mass increase should be 2.5%.
The same model gives the influence of sweep on wing mass as
$$m_{Wing}\propto\frac{1}{cos\varphi^{\,0.57}}$$
Based on 40° wing sweep, every decrease in spar thickness of 1% requires a sweep decrease of about 2.9° in order to keep wing weight constant, based on that relation. If that 1% thickness decrease allows you to desweep the wing by 3° or more, you come out ahead.
As always, a compromise is the best solution. Airliners of the propeller era had thicker wings than jet airliners. Some sweep and some thickness reduction will allow to push the critical Mach number to the highest value. Tank volume is indeed an important factor and is responsible for the comparatively low aspect ratios of early jet airliners.

Does the exponent of the second equation also change?

@Abdullah All these exponents depend heavily on the composition of the equation. When - for example - dive speed is part of the mix, it helps to "hide" the influence of sweep since both are closely correlated. So the exponent will already change between equations (Roskam - Douglas - Boeing - Datcom - LTH, and I'm sure the Russians have their own version, too) and equally when a different mix of planes is used.

Also, you have to think that the air does not hit the wing at a 90* angle to the leading edge on a swept wing.  The shape of the wing ribs is not necessarily the shape that the air is being in contact with.  It is probably going to be a longer trip than the shape of the rib in certain areas of the wing, like close to the tip, where both leading and trailing edges have some sweep to them.


text:
Is the ideal airfoil concave on the underside?
Virtually all airplane wings are convex on both the top and at least part of the bottom. But thus convexity on the bottom creates suction which reduces lift. I can only think of two reasons to have it: negative AoA performance (important only for aerobatics), and the need to accommodate a spar. Suppose these two were possible to neglect. Would a concave bottom give a better L/D?
So, is the L/D a cruising aircraft best when bottom is concave?
When you look at applications where the structural and AoA concerns are absent, it does seem that the answer is yes.

The blackbird is small, hence there are less structural concerns. It has a concave bottom.
Turbofan blades are also relatively small. Also, I think the forces are dominated by centrifugal force, not bending force. Once again, the bottom.is concave.
The same is true of the turbine blade.
Curiously, the propeller, which also has similar conditions, has a convex bottom.

There is no ideal airfoil for all situations - but optimized for the specific szenario it is designed to operate in - and in most cases, there will be quite a lot of compromises...

Get the book The Miracle of Flight by Stephen Dalton. It's written for laypeople. He goes though aerodynamics and lift generation, and mechanics, starting with insects. then birds, and moving up to aircraft. At very low Reynolds Numbers, heavily cambered single surface wings work best. As RN goes up, wings get thicker until the thickness becomes a drag problem at high speed, then they get thinner again. The ideal supersonic airfoil is a flat thin uncambered sheet.  Efficiency of undercamber depends on the operating angle of attack.

So, is the L/D [of] a cruising aircraft best when bottom is concave?

If we use gliders as a reference, the bottom of the best airfoils is only very slightly concave, almost straight, in order to allow for a long laminar run while still achieving high lift. Up to the Nineties of the last century, the best designs used rear loading over the last 15% - 20% while the most recent ones do away with even that to squeeze a few percent more of laminarity from the lower side. This has to do with flying technique as well as better understanding of laminar flow: Instead of optimizing the design for two polar points (thermalling and fast flight), now the ideal is to also look at intermediate speeds like when flying under lined-up clouds.
While birds have wings optimized for low speed because they can retract and fold their wings when going fast, airplanes have to design their wings for high speed and add flaps and slats for low speed flight. When you look at an airfoil of an airliner in landing configuration, it is not so different from a bird's wing.

Typical landing configuration of an airliner wing, from an article by A. M. O. Smith, McDonnell-Douglas, in Journal of Aircraft, Vol 12 No 6, 1975.
Also, birds do not have to deal with trans- and supersonic effects which dictate the airfoil shape of almost all jet-powered airplanes.

Curiously, the propeller, which also has similar conditions, has a convex bottom.

Propeller and especially rotor airfoils should have a low pitching moment change over angle of attack to avoid twisting the propeller rsp. rotor blade. This is achieved by reduced camber, and the result is a convex bottom surface: A symmetrical airfoil has no pitching moment variation within the linear range of angles of attack. In order to be able to work over a range of angles of attack, they must have a certain thickness, which of course also benefits the structure.

Is the parasitic drag less for the cambered "low speed" airfoil?

@AbdullahisnotanAmalekite Depends on angle of attack. Generally, thinner airfoils have less parasitic drag over a small angle of attack range while thicker airfoils have a bit more parasitic drag over a wider angle of attack range.

What about the lowest achievable values?

@AbdullahisnotanAmalekite Flat plate at zero angle of attack, obviously. Equals twice the friction drag at the respective Reynolds number. Another interesting concept uses a slotted flap for pressure recovery to enable laminar flow over the first 80% of both sides. See here for more.

Ooof. Ok, let's rephrase it. Which one can achieve the higher lift to parasitic drag ratio?

@AbdullahisnotanAmalekite That would be a thin, birdlike airfoil like a cambered plate similar to the Göttingen 417.

Ironic, given how similar it is to the early biplane airfoils, and how dissmilar it is to almost anything modern

@AbdullahisnotanAmalekite Early biplanes essentially flew at one speed and with little mass variation, so this made sense if selected properly. However, all the wires destroyed any parasitic drag advantage. It is better to use a thicker airfoil in a cantilever wing.

Back to the roots. An airfoil can be basically described by 3 parameters:

the shape/curvature of its camber line;
the radius of its leading edge (the trailing edge is almost always pointy);
the distribution of the thickness upon the camber line.

Those 3 parameters are enough to define any airfoil of practical use (source Wikipedia):

Without any claim to thoroughness (that would cover an entire book) the following can be said:

a rounder leading edge and higher thickness give a smooth stall with a progressive loss of lift; pitching moment also changes smoothly approaching the stall region; thick airfoils have also a big volume that can be used for fuel and structure $\rightarrow$ good for the root of a jetliner's wing or of a propeller.
a thinner leading edge and thickness give a more nervous stall with a higher pitching moment variation; anyway at high subsonic speeds a thinner airfoil possesses a lower drag than a thick one since it retards the formation of shock waves (which give rise to the boundary layer's detachment and stall) $\rightarrow$ due to structural and aerodynamical reasons, lift on a jetliner's wing and on a propeller should diminish toward the tip: twisting the wing/propeller so that the AoA at its tip is lower than that at the root allows the use of these thinner airfoils with a benefit in terms of drag reduction.
a pointy leading edge is necessary at supersonic speeds in order to avoid the formation of a detached (aka bow) shock which increases drag and temperature $\rightarrow$ virtually all the wings of supersonic fighters are built around a very thin, pointy airfoil; the loss in internal volume for fuel and structure is compensated for by the use of delta planform for the wing.
camber is mainly used to control/delay stall and limit pitching moment; a very pronounced camber (one that makes the airfoil look like an arc) is used when lift needs to be generated at very high AoA i.e. low speeds; this is typically the shape that a bird's wing possesses or that a jetliner's wing assumes when its high-lift devices are deployed; in those cases a round, thick leading edge helps too, as already said; anyway a lot of camber doesn't work at high speeds since the boundary layer tends to detach and make the airfoil stall; the only exception where high camber and high speed work together is within an axial compressor/turbine: here the adjacent blades "channel" the airflow and make it stay attached also at the high speeds seen by the blades. As a side effect, an airfoil with high camber generates also a high pitching moment which needs a lot of structure to be carried: also for this reason the turbine blade is quite tick; to partially offset this, the camber can be built in such a way to go upward toward the trailing edge: so the first, say, ¾ of the chord is arched like the turbofan blade in your picture while the last ¼ is arched the other way around, in a sort of "S" shaped pattern; in this way the good aerodynamic characteristics at high AoA are retained but without the increased torsional moment $\rightarrow$ this is normally seen in helicopter blades (CH-47, source)

To get to your original question:

Is the ideal airfoil concave on the underside?

A highly cambered airfoil (such that also the belly is arched) works only when lift is needed at high AoA and low speeds and the high torsional (aka pitching) moment doesn't play an important role. For airfoils in cascade (like in an axial compressor/turbine) the limitation on the speed can be dropped due to the channeling effect. For all the other cases a more subtle camber or even an S-shaped camber is needed.

is the L/D for a cruising aircraft best when the bottom is concave?

It depends on the Reynolds number.  Full scale aircraft have a high enough value that they can create most of their required lift with the top of the wing. Even though the bottom of the wing is convex, it still contributes to lift due to its positive angle of attack.
In cruise, AoA will be small, and the convex wing bottom best serves by being a very low drag shape and adding to wing structural strength.
"Top lift" provides the best L/D up to the airspeed where Mach effects increase drag of the upper wing.
The effects of Reynolds number on L/D can be seen at airfoil tools website.
Cambering the underside does provide a higher coefficient of lift, but at the expense of higher drag.  Where efficiency is not important, such as landing, adding camber (with flaps) can be very useful in lowering landing speed.


text:
How do electric anti-icing systems work?
How do they work for GA propeller planes? More specifically, how is it implemented on the leading edge of the wing and propeller blades? I'm more interested in the mechanics of electric heating, not the theory or where it gets power from (alternator or batteries).
What are the pros and cons of electric anti-icing compared to other systems?

If they weren't viable they wouldn't be used @JoshFang

This is a very broad question. Aircraft electric anti-icing systems including air data sensor (temperature and pressure probes, e.g. pitot probe), windscreens, propellers, and wings. Each system has a different design and implementation. The direct answer to the question as asked is: by using DC or AC current resistance to heat a material; each system is implemented according to the design requirements of that system; and they are viable. This question should be narrowed to a specific design or subset of designs.

Thanks for the response. Edited to narrow down the question.

How do they work?
An electric current is passed through a heating element built into the device needed to be heated. For example, 115 VAC may be applied to a heating element of anywhere between 5-300 ohms (resistance values vary widely depending on the heating application). The voltage applied may be dependent on a controller that will dictate how much voltage is applied.
Using basic Ohm's Law, we can deduce that the current draw of 115 VAC across an example 60 ohms coil, would be approximately 1.9 A, and the power 220W. Keep in mind, this isn't a strict calculation as phase angle and other variables might be involved in an AC circuit - but close enough... That 220W is dissipated as heat which thus heats the device.
This is virtually the same in most GA and commercial applications.
Are they viable?
If by viable you mean "They work, but do have occasional failures that falls within a reasonable rate of expectation," then yes, they are viable.

Small GA aircraft have made use of electric anti-ice systems on the wing and tailplanes.  The most common one being ThermaWing, which consists of electrically heated laminate adhered to the leading edges of the wings and tailplane.  They are powered by a separate alternator driven by the engine.  They have an advantage over chemical deicing methods like TKS in that they are not time limited by the quantity of deicing chemical which can be carried at the cost of a lowered useful load due to the weight of the additional alternator plus the wing cuffs and other equipment.  As to the best of my knowledge ThermaWing is an STC aftermarket kit and does not provide certified Flight Into Known Icing (FIKI) unlike some TKS applications. Columbia aircraft used to offer ThermaWing as optional on their 350 and 400 series aircraft prior to their bankruptcy and acquisition by Cessna.
Electric propeller deicing - hot props - consist of a rubber sleeve applied to the inboard most sections of the propller blade leading edge, each one containing an electric heating element underneath which could melt accumulated ice.  Again it competes directly with chemical deicing systems which sling deicing fluid over the length of the propeller blade to melt ice and has similar benefits and drawbacks to the ThermaWing system.
Windshield deicing and defogging is usually done with cabin heat air coming from either outside air passed over the exhaust manifold then vented into the cabin or from a dedicated fuel burning cabin heater, common on twins.  Still some light aircraft eg the BE58 Baron make use of electriccally heated windshield sections consisting of a finely woven heating element mesh applied in the transparency laminate and can be powered by the airplane’s electrical system.
Pitot-static ports, stall warning vanes, and fuel vents also often make use of electric heating elements to prevent accretion of ice which could block these ports and cause primary flight instruments to malfunction or cause fuel starvation of the engines.

all_licenses: Creative Commons - Attribution Share-Alike - https://creativecommons.org/licenses/by-sa/3.0/


text:
Are there any models for planning aircraft placement in hangars?
I work alone at night doing line operations (fueling, tugging/towing, etc.) and emergency services (police, fire and EMS) at my airport. When a pilot/customer comes in and wants to know if we have hangar space available for their aircraft, it would be nice when planning aircraft placement in those hangars to have figures to scale of various aircraft. That way you can see whether or not a specific aircraft can indeed fit, and in what configuration, before trying as I rarely (if ever) have a "spotter" or a wing-walker.
Or perhaps even software or a computer program. Put in the hangar dimensions and which aircraft are to be placed.

I guess it would be the same as the software used to cut shapes out of (one-sided) fabric, while minimising waste, but I like the model idea - really quick and intuitive.

https://staxplanes.com/ appears to do what you want.

I’ve not heard of software programs to do this, but the Navy has made use of a relatively low tech solution to address this problem for decades aboard their aircraft carriers.

Known simply as the Ouija Board, the aircraft handler and his crew utilize a scaled physical map of both the carrier flight deck and hangar deck along with appropriately scaled footprint models of aircraft in order to track and plan aircraft movements.
I’d guess a similar scheme could be employed to solve your problems.

You could use magnets on the backs of the footprint models to make a version that hangs on a wall, for that matter, if horizontal space was at a premium

This is cool, but you need accurate scale models.  This is much easier on a carrier, where there aren't many different models of plane, than in an airport situation like the OP describes, I'd imagine.  "Can you fit my plane in your hangar"?  "IDK, did you bring a 1/32nd scale model of it?"

If there is no algorithm , then using a painting software can do it too. Or a developer can write 50 lines of Javascript.

There’s at least one software solution that can be used to address this for a wide range of aircraft types. In my old company, we have used a software called PathPlanner and looked at another one called Aeroturn to simulate aircraft ground manoeuvring. The same or similar software products are also used for airport design. They would, as far as I can see, very well allow to simulate hangar occupancy as well.

We solved a similar problem of optimization for a customer once which is an analog of this.  The exact algorithm used was simulated annealing.  If the same problem were presented today, there would be many competing alternative algorithms, generally in the class of genetic algorithms.
The problem of optimization presents itself in many aviation related activities, from packing aircraft, to routing aircraft, and in a recent instance, to optimizing dynamically flights in a battlefield, to factor in new situations, and adjust the costs and risks depending upon the rules established by command.
The interested person might start with an analog of the problem, which is the traveling salesman problem. One might also consider watching: "The Secret Rules of Modern Living: Algorithms," a one hour movie on similar topics, where air traffic sequencing is discussed.
In short, the problem is trivial with a handful of planes to fit in a hangar, and grows rapidly as the numbers increase.  So much so that it rapidly becomes seemingly unsolvable.  Optimization algorithms generally do not find perfect solutions, but find approximations, which are practically close enough.

A former client disclosed to me that they solve this problem using a interference detection software package running on AutoCAD.

There is a program called Stax that runs on Mac, Windows, and iPad that probably does what you're looking for.

Is there a link to the specific feature on their site? I can't seem to find it mentioned, but the site is kind of a mess.

staxplanes.com is the website


text:
What is the difference between climb angle and flight path climb angle?
Consequently, which of these two angles does wind affect?
Thank you everyone!

What makes you think they're different at all?

You possibly want to know the difference between the pitch angle of the aircraft (attitude) and the flight path angle (climb angle).

Maybe this is about the difference between „aerodynamic climb angle“ and flight path angle, the former being ARCSIN(VS/TAS) and the other one ARCTAN(VS/GS); with VS vertical speed, TAS true air speed, GS ground speed. The former would depend only on aircraft climb performance, the latter additionally on wind.

If you climb a flight of stairs on a cruise ship traveling at 50 kts, does the speed of the cruise ship affect how steep the stairs are?

@CharlesBretana: When something is moving, you therefore have two frames of reference, the "fixed" one, and the moving one. When you talk about ship speed and stair dimensions, you use different frames, the "steepness" of the stairs in the user frame of reference is unchanged, regardless of the ship speed, as long as the acceleration is null. You could also introduce the orbital speed of Earth (100,000 km/h), and the Milky Way drift, that wouldn't change anything for the user.

@mins,  actually, the "steepness" of the stairs, like  everything else, is totally dependent on what frame of reference you measure it in. That's the entire point of my comment. Too often, we lose sight of the simple "change of frame of reference" and the Transformation necessary for any and all measurements made in one frame of an event or property in another frame. If it takes you a minute to climb the stairs, and the cruise ship travels 100 feet/minute, then you have climbed vertically the height of the stairs, and horizontally the length of the stairs plus 100 feet.

The point is that the issue ALWAYS depends on which frame of reference you measure in. And all frames are equally valid.  Everything is relative.  The op's question is about nothing more than understanding this. The "Flight Path angle" (measured in the frame of the moving block of air), is different than the "climb angle" measured in the frame of reference of the ground. Which one of these the phrase "climb gradient" refers to is only definitional, (semantics). It is not substantive.

@CharlesBretana: Ok, I got what you meant and wanted to convey.

@mins,  No worries. I understand. (although, respectfully, I might point out that even using the language as you did, referring to the two frames of reference as "fixed", and "moving" implies that one of them, the "fixed" one, is somehow more "real" than the other, and this implication, or understanding that some might get from reading this language, is the crux of the issue. No frame of reference is any more "real" than any other.

@CharlesBretana: That's the reason "fixed" is between quotes, contrary to "moving".

ahhh, I see your intent now. Sorry.


text:
What is the up-to-date disposition towards narcolepsy in the U.S.A. for obtaining a pilot license?
Citation: What medical conditions does the FAA consider disqualifying?
In the past, I recall this page having listed narcolepsy, but I find no mention of narcolepsy anywhere relevant on the FAA website anymore. Can anyone enlighten me as to whether I am just missing it somewhere, or if there is knowledge of this specifically having been changed? It would surprise me greatly if this were not a disqualifying condition, but I want to find official information on it.

I would hope that a condition that makes you suddenly fall asleep out of the blue would be disqualifying...

@Sean that's actually a misconception about narcolepsy- while possible, lots of people have much more mild forms of narcolepsy that only manifest as daytime drowsiness.

Which is still extremely dangerous in a pilot.

In the FAA Guide For Medical Examiners there is only a single direct reference to narcolepsy
Under the Obstructive Sleep Apnea QA section it states

Does this process involve other sleep disorder conditions? (E.g.
  Period Limb Movement Disorder, narcolepsy, central sleep apnea, etc.)?
No. This process is for obstructive sleep apnea only. If it is clear
  that the airman suffers from a different sleep disorder, DEFER and
  submit any supporting documentation for FAA decision.

This would indicate that if you suffer from any sleep disorder conditions your medical should be DEFERed
Earlier in the document it states:

Medical conditions that chronically interfere with sleep are
  disqualifying regardless of whether a sleep aid is used or not.
  Examples may include primary sleep disorders (e.g., insomnia, sleep
  apnea) or psychological disorders (e.g., anxiety, depression). While
  sleep aids may be appropriate and effective for short term symptomatic
  relief, the primary concern should be the diagnosis, treatment, and
  resolution of the underlying condition before clearance for aviation
  duties.

Narcolepsy is a sleep disorder and would potentially fall under this definition and thus prevent your from getting clearance.
In these cases its usually best to get the advice of an FAA AME.

Thank you for this very direct and cited information. I really appreciate it.